As shown in FIG. 1, there is shown a cross-section through a portion of a combustion turbine 10 may include a compressor section 12, a combustion section 14 and a turbine section 16. A rotor assembly 18 is centrally located and extends through the three sections. The compressor section 12 may include cylinders 20, 22 that enclose alternating rows of stationary vanes 24 and rotating blades 26. The stationary vanes 24 may be affixed to the cylinder 20 while the rotating blades 26 may be mounted to the rotor assembly 18 for rotation with the rotor assembly 18.
The combustion section 14 may include a shell 28 that forms a chamber 30. Multiple combustors, for example, sixteen combustors (only one combustor 32 of which is shown) may be contained within the combustion section chamber 30 and distributed around a circle in an annular pattern. Fuel 34, which may be in liquid or gaseous form, such as oil or gas, may enter each combustor 32 and be combined with compressed air introduced into the combustor 32 from the chamber 30, as indicated by the unnumbered arrows surrounding the combustor 32. The combined fuel/air mixture may be burned in the combustor 32 and the resulting hot, combustor exhaust gas flow 36 may be exhausted to a transition section 38 attached to the combustor 32 for routing to the turbine section 16.
The turbine section 16 may include a cylindrical housing 40, including an inner cylinder 42, and may enclose rows of stationary vanes and rotating blades, including vanes 44 and blades 46. The stationary vanes 44 may be affixed to the inner cylinder 42, and the rotating blades 46 may be affixed to discs that form parts of the rotor assembly 18 in the region of the turbine section 16. The first row of vanes 44 and the first row of blades 46 near the entry of the turbine section 16 are generally referred to as the first stage vanes and the first stage blades, respectively.
Encircling the rotor assembly 18 in the turbine section 16 may be a series of vane platforms 48, which together with rotor discs 50, collectively define an inner boundary for a gas flow path 52 through the first stage of the turbine section 16. Each transition section 38 in the combustion section 14 may be mounted to the turbine section housing 40 and the vane platforms 48 to discharge the gas flow 30 towards the first stage vanes 44 and first stage blades 46.
In operation, the compressor section 12 receives air through an intake (not shown) and compresses it. The compressed air enters the chamber 30 in the combustion section 14 and is distributed to each of the combustors 32. In each combustor 32, the fuel 34 and compressed air is mixed and burned. The hot, combustor exhaust gas flow 36 is then routed through the transition section 38 to the turbine section 16. In the turbine section 16, the hot, compressed gas flow is turned by the vanes, such as first stage vane 44, and rotates the blades, such as first stage blade 52, which in turn drive the rotor assembly 18. The gas flow is then exhausted from the turbine section 16. The turbine system 10 may include additional exhaust structure (not shown) downstream of the turbine section 16. The power thus imparted to the rotor assembly 18 may be used not only to rotate the compressor section blades 26 but also to rotate other machinery, such as an external electric generator or a fan for aircraft propulsion (not shown).
Referring now to FIG. 2, three adjacent transition sections 38 are shown as when viewed from axially downstream. Although the transition sections are substantially identical, the transition section 38 located at the 12 o'clock position is used to discuss the relevant parts of each. It should be understood that a turbine engine would have additional transition sections, for example, a total of sixteen, spaced in an annular array.
The transition section 38 may include a transition section body 56 having an inlet 58 for receiving a gas flow exhausted by an associated combustor (not shown, but see FIG. 1). The transition section body 56 may include an internal passage 60 from the inlet 58 to an outlet 62 from which the combustor exhaust gas flow 36 is discharged towards the turbine section (not shown). Surrounding the transition section body 56 is the compression chamber 30, which contains compressed gas that has not yet entered a combustor.
FIG. 3 shows parts of three adjacent transition sections 38 in cross section. Each transition section 38 includes a transition section body 56, an outlet 62, and a combustor exhaust gas flow 36. Surrounding each transition section 38 is the compression chamber 30 containing compressed gas that has yet to enter a combustor (not shown). Downstream of the outlet 62, the combustor exhaust flow 36 enters the turbine entry zone 64, which is the turbine section upstream of the first row of vanes or blades (not shown). In the turbine entry zone 64, the combustor exhaust flow 36 mixes with leaked compressor gas 66 that passed through a gap 68 between adjacent transition sections 38. This leaked compressor gas 66 bypasses the combustors (not shown) and reduces efficiency of the combustion turbine (see FIG. 1). This leakage is driven by the pressure drop between the higher pressure air in the compression chamber 30 and the lower pressure compressor gas in the turbine entry zone 64.
Excess leakage through the gaps between adjacent transitions may prevent a combustion turbine from achieving optimal performance. The power generated by a combustion turbine is, in large part, a function of the initial temperature of the gas expanded through the turbine section. Because the efficiency of a combined cycle turbine process depends on the turbine inlet temperature, the higher the temperature of the gas entering the turbine through the turbine entry zone the more efficient the combined cycle turbine process. The temperature of the gases in the turbine entry zone is referred to as the thermodynamically relevant process temperature.
Compressed air that leaks into the turbine entry zone without passing through the combustors reduces the thermodynamically relevant process temperature because the compressed air is significantly cooler than the combustor exhaust gas flow. Similarly, if all the compressed air in the combustion chamber is fed directly through the combustors, the thermodynamically relevant process temperature would be increased while maintaining a constant combustion temperature. Thus, improved seals for the gap between adjacent transition sections would help improve performance for a combustion turbine.
Controlling or preventing leakage through the gaps between adjacent transition sections is complicated by a number of factors. For instance, differences in thermal expansion cause the size of the gap to change during the operational cycle of the combustion turbine. Another factor is thermal distortion due to temperature gradients that may cause the size of the gap to change along the length of the gap. Another factor is the pressure drop between the compression chamber gases and the turbine entry zone gases, which may typically be 0.5 bar or more. Finally, another factor is the temperature in the gap between adjacent transition sections reaches temperatures ranging from 500° C. to 800° C. during operation of the combustion turbine.
Currently, labyrinth seals and brush seals have been used to limit leakage through the gap between adjacent transition sections. However, labyrinth seals can only achieve limited sealing effectiveness within the available axial space, which is limited by the thickness of the transition section body. In addition, brush seals are comparatively expensive, show limited life, and lead to increased life cycle costs.
Accordingly, a need exists for an improved seal that controls gas leakage across a gap with a high pressure drop, such as the gap between adjacent transition sections that leads from the compression chamber to the turbine entry zone.